Method for treating a gas turbine blade and gas turbine having said blade

ABSTRACT

The use of different ceramic layers allows different configurations of gas turbines to be produced each of which is optimized for a respective use of base load operation or peak load operation.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a 35 U.S.C. §§371 national phase conversionof PCT/EP2012/069700, filed Oct. 5, 2012, the contents of which areincorporated by reference herein. The PCT International Application waspublished in the German language.

TECHNICAL FIELD

The invention relates to a process for producing gas turbines offlexible design, to gas turbines and to methods for operating gasturbines.

TECHNICAL BACKGROUND

For generating electricity, gas turbines can be operated in base loadoperation or in particular in peak load operation. The demands forsatisfaction of these respective conditions are different. An optimizedconfiguration of the gas turbine which satisfies both demands wouldalways represent a compromise. It is therefore an object of theinvention to solve this problem.

BRIEF DESCRIPTION OF THE DRAWINGS

FIGS. 1-3 show an exemplary embodiment of the invention,

FIG. 4 shows a pore distribution of a ceramic coating,

FIG. 5 shows a turbine blade or vane, and

FIG. 6 shows a gas turbine.

DESCRIPTION OF AN EMBODIMENT

The description represents merely an exemplary embodiment of theinvention.

A maintenance interval for gas turbine 100 (which is shown in FIG. 6) isdetermined by ascertaining the operational hours and starts, and theseare dependent on the mode of operation and specific factors. Themaintenance is to be carried out in each case when the hour or startlimit has been reached.

Depending on the field of use of the gas turbine, if it is thennecessary to carry out maintenance or if the use demands reconditioningor another use beforehand, the configuration of the gas turbine 100 isaltered.

DEFINITIONS OF THE TERMS

First gas turbine has first turbine blade or vane with first thermalbarrier coating.

Second gas turbine has turbine blades or vanes with ceramic thermalbarrier coatings,

a) in which the first turbine blades or vanes (=second turbine blade orvane) and/orb) new, unconsumed turbine blades or vanes (=new, second turbine bladesor vanes)are used,and in each case have a second thermal barrier coating which can beclearly distinguished from the first thermal barrier coating.

If a single-layer thermal barrier coating was present in operationbeforehand in said first gas turbine, as described above, a two-layer(FIG. 3), a thicker (FIG. 1) or a more porous ceramic thermal barriercoating is used for the turbine blades or vanes 120, 130 for the reneweduse in base load operation.

The origin (that is, the same substrate) of the turbine blades or vanesfor the second gas turbine [[can]] may be the first turbine blades orvanes of the first gas turbine or other gas turbines, which were alreadyin use, have been appropriately refurbished and then give rise to secondturbine blades or vanes through recoating, or can be new, second turbineblades or vanes, in which newly produced (newly cast) turbine blades orvanes which have not yet been used are coated differently to the firstturbine blades or vanes of the first gas turbine.

Similarly, it is possible, if the gas turbine 100 had a two-layerceramic thermal barrier coating on the turbine blades or vanes 120, 130in base load operation, to apply a single-layer TBC, such that it canthen be used in peak load operation (daily starter) (FIG. 2).

For peak load operation, it is preferable to use only a single-layerceramic coating with a uniform porosity. For peak load operation, theceramic thermal barrier coating on the turbine blades or vanes 120, 130preferably has a high porosity of 18%±4%.

In base load operation (base loader), however, a two-layer thermalbarrier coating 13 is used (FIG. 3).

It is preferable to use agglomerated, sintered powder as starting powderfor the ceramic coatings 7′, 7″, 7′″, 10′, 13′.

Each ceramic sprayed coating is applied in coating layers. Two-layernature means, however, that a second layer differs from a first,underlying layer in terms of porosity and/or microstructure and/orchemical composition.

A ceramic layer 7 with a porosity of 12%±4% which preferably has acoating thickness of 75 μm to 150 μm is preferably used as the bottomlayer.

A layer with a porosity of 25%±4% is sprayed or is present thereabove asthe outer ceramic layer 10.

The difference in the porosity is, however, at least 2%, in particularat least 4%. Variations in the porosity during production are known. Novariations are to be recorded within a charge, i.e. a blade or vane set.

In order to generate porosities in ceramic coatings or ceramic layers(FIGS. 1-3), the spraying can involve the use of coarse grains and usecan be made of polymers or smaller grains with polymer, coarse meaningan at least 20% greater mean particle diameter.

A two-layer ceramic coating 7, 10 can be produced using differentspraying processes: the bottom layer 7 is sprayed without polymer andthe top layer 10 is sprayed with polymer.

This gives rise to larger pores in the top layer 10, i.e. the mean porediameter d₁₀ increases compared to the mean pore diameter d₇ of thebottom layer 7 (FIG. 4). This is not necessarily the case. A higherporosity is often only achieved by a higher number of pores of the samepore size.

It is preferable that the same powder is used in this case, i.e. also anidentical grain size distribution.

Zirconium oxide (ZrO₂) for the ceramic layers of the thermal barriercoatings preferably has a monoclinic proportion of 3%, in particular≦1.5%. A ceramic layer or coating 7, 7′, 10, 13 (FIGS. 1-3) on theturbine blade or vane 120, 130 then has corresponding proportions.

The minimum proportion of the monoclinic phase is at least 1%, inparticular 0.5%, so as not to excessively increase the costs of thepowder.

The change in the configuration of the first thermal barrier coating 7′,7″, 13′ virtually produces another, second gas turbine optimized for itsfield of use.

FIG. 5 shows a perspective view of a rotor blade 120 or guide vane 130of a turbomachine, which extends along a longitudinal axis 121.

The turbomachine may be a gas turbine of an aircraft or of a power plantfor generating electricity, a steam turbine or a compressor.

The blade or vane 120, 130 has, in succession along the longitudinalaxis 121, a securing region 400, an adjoining blade or vane platform403, a main blade or vane part 406 and a blade or vane tip 415.

As a guide vane 130, the vane 130 may have a further platform (notshown) at its vane tip 415.

A blade or vane root 183, which is used to secure the rotor blades 120,130 to a shaft or a disk (not shown), is formed in the securing region400.

The blade or vane root 183 is designed, for example, in hammerhead form.Other configurations, such as a fir-tree or dovetail root, are possible.

The blade or vane 120, 130 has a leading edge 409 and a trailing edge412 for a medium which flows past the main blade or vane part 406.

In the case of conventional blades or vanes 120, 130, by way of examplesolid metallic materials, in particular superalloys, are used in allregions 400, 403, 406 of the blade or vane 120, 130.

Superalloys of this type are known, for example, from EP 1 204 776 B1,EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The blade or vane 120, 130 may in this case be produced by a castingprocess, also by means of directional solidification, by a forgingprocess, by a milling process or combinations thereof.

Workpieces with a single-crystal structure or structures are used ascomponents for machines which, in operation, are exposed to highmechanical, thermal and/or chemical stresses. Single-crystal workpiecesof this type are produced, for example, by directional solidificationfrom the melt. This involves casting processes in which the liquidmetallic alloy solidifies to form the single-crystal structure, i.e. thesingle-crystal workpiece, or solidifies directionally.

In this case, dendritic crystals are oriented along the direction ofheat flow and form either a columnar crystalline grain structure (i.e.grains which run over the entire length of the workpiece and arereferred to here, in accordance with the language customarily used, asdirectionally solidified) or a single-crystal structure, i.e. the entireworkpiece consists of one single crystal. In these processes, atransition to globular (polycrystalline) solidification needs to beavoided, since non-directional growth inevitably forms transverse andlongitudinal grain boundaries, which negate the favorable properties ofthe directionally solidified or single-crystal component.

Where the text refers in general terms to directionally solidifiedmicrostructures, this is to be understood as meaning both singlecrystals, which do not have any grain boundaries or at most havesmall-angle grain boundaries, and columnar crystal structures, which dohave grain boundaries running in the longitudinal direction but do nothave any transverse grain boundaries. This second form of crystallinestructures is also described as directionally solidified microstructures(directionally solidified structures).

Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0892 090 A1.

The blades or vanes 120, 130 may likewise have coatings protectingagainst corrosion or oxidation, e.g. (MCrAlX; M is at least one elementselected from the group consisting of iron (Fe), cobalt (Co), nickel(Ni), X is an active element and stands for yttrium (Y) and/or siliconand/or at least one rare earth element, or hafnium (Hf)). Alloys of thistype are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 orEP 1 306 454 A1.

The density is preferably 95% of the theoretical density. A protectivealuminum oxide layer (TGO=thermally grown oxide layer) is formed on theMCrAlX layer (as an intermediate layer or as the outermost layer).

The layer preferably has a composition Co-30Ni-28Cr-8Al-0.6Y-0.7Si orCo-28Ni-24Cr-10Al-0.6Y. In addition to these cobalt-based protectivecoatings, it is also preferable to use nickel-based protective layers,such as Ni-10Cr-12Al-0.6Y-3Re or Ni-12Co-21Cr-11Al-0.4Y-2Re orNi-25Co-17Cr-10Al-0.4Y-1.5Re.

It is also possible for a thermal barrier coating, which is preferablythe outermost layer and consists for example of ZrO₂, Y₂O₃—ZrO₂, i.e.unstabilized, partially stabilized or fully stabilized by yttrium oxideand/or calcium oxide and/or magnesium oxide, to be present on theMCrAlX.

The thermal barrier coating covers the entire MCrAlX layer. Columnargrains are produced in the thermal barrier coating by suitable coatingprocesses, such as for example electron beam physical vapor deposition(EB-PVD).

Other coating processes are possible, for example atmospheric plasmaspraying (APS), LPPS, VPS or CVD. The thermal barrier coating mayinclude grains that are porous or have micro-cracks or macro-cracks, inorder to improve the resistance to thermal shocks. The thermal barriercoating is therefore preferably more porous than the MCrAlX layer.

Refurbishment means that after they have been used, protective layersmay have to be removed from components 120, 130 (e.g. by sand-blasting).Then, the corrosion and/or oxidation layers and products are removed. Ifappropriate, cracks in the component 120, 130 are also repaired. This isfollowed by recoating of the component 120, 130, after which thecomponent 120, 130 can be reused.

The blade or vane 120, 130 may be hollow or solid in form. If the bladeor vane 120, 130 is to be cooled, it is hollow and may also havefilm-cooling holes 418 (indicated by dashed lines).

FIG. 6 shows, by way of example, a partial longitudinal section througha gas turbine 100.

In the interior, the gas turbine 100 has a rotor 103 with a shaft 101which is mounted such that it can rotate about an axis of rotation 102and is also referred to as the turbine rotor.

An intake housing 104, a compressor 105, a, for example, toroidalcombustion chamber 110, in particular an annular combustion chamber,with a plurality of coaxially arranged burners 107, a turbine 108 andthe exhaust-gas housing 109 follow one another along the rotor 103.

The annular combustion chamber 110 is in communication with a, forexample, annular hot-gas passage 111, where, by way of example, foursuccessive turbine stages 112 form the turbine 108.

Each turbine stage 112 is formed, for example, from two blade or vanerings. As seen in the direction of flow of a working medium 113, in thehot-gas passage 111 a row of guide vanes 115 is followed by a row 125formed from rotor blades 120.

The guide vanes 130 are secured to an inner housing 138 of a stator 143,whereas the rotor blades 120 of a row 125 are fitted to the rotor 103for example by means of a turbine disk 133.

A generator (not shown) is coupled to the rotor 103.

While the gas turbine 100 is operating, the compressor 105 sucks in air135 through the intake housing 104 and compresses it. The compressed airprovided at the turbine-side end of the compressor 105 is passed to theburners 107, where it is mixed with a fuel. The mix is then burnt in thecombustion chamber 110, forming the working medium 113. From there, theworking medium 113 flows along the hot-gas passage 111 past the guidevanes 130 and the rotor blades 120. The working medium 113 is expandedat the rotor blades 120, transferring its momentum, so that the rotorblades 120 drive the rotor 103 and the latter in turn drives thegenerator coupled to it.

While the gas turbine 100 is operating, the components which are exposedto the hot working medium 113 are subject to thermal stresses. The guidevanes 130 and rotor blades 120 of the first turbine stage 112, as seenin the direction of flow of the working medium 113, together with theheat shield elements which line the annular combustion chamber 110, aresubject to the highest thermal stresses.

To be able to withstand the temperatures which prevail there, they maybe cooled by means of a coolant.

Substrates of the components may likewise have a directional structure,i.e. they are in single-crystal form (SX structure) or have onlylongitudinally oriented grains (DS structure).

By way of example, iron-based, nickel-based or cobalt-based superalloysare used as material for the components, in particular for the turbineblade or vane 120, 130 and components of the combustion chamber 110.

Superalloys of this type are known, for example, from EP 1 204 776 B1,EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The blades or vanes 120, 130 may likewise have coatings protectingagainst corrosion (MCrAlX; M is at least one element selected from thegroup consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an activeelement and stands for yttrium (Y) and/or silicon, scandium (Sc) and/orat least one rare earth element, or hafnium). Alloys of this type areknown from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306454 A1.

A thermal barrier coating, consisting for example of ZrO₂, Y₂O₃—ZrO₂,i.e. unstabilized, partially stabilized or fully stabilized by yttriumoxide and/or calcium oxide and/or magnesium oxide, may also be presenton the MCrAlX.

Columnar grains are produced in the thermal barrier coating by suitablecoating processes, such as for example electron beam physical vapordeposition (EB-PVD). The guide vane 130 has a guide vane root (not shownhere), which faces the inner housing 138 of the turbine 108, and a guidevane head which is at the opposite end from the guide vane root. Theguide vane head faces the rotor 103 and is fixed to a securing ring 140of the stator 143.

1. A process for producing a second gas turbine which has turbine bladesor vanes with respective ceramic thermal barrier coatings, the methodcomprising: from a first gas turbine which has first turbine blades orvanes with respective thermal barrier coatings, removing at least onefirst ceramic thermal barrier coating from the first turbine blades orvanes of the first gas turbine; at least one of applying a new, secondceramic thermal barrier coating to the first turbine blades or vanesfrom which the coating has been removed to produce second turbine bladesor vanes, or applying a new, second ceramic thermal barrier coating tonew, second turbine blades or vanes which have been newly produced;wherein the second ceramic thermal barrier coatings differ significantlyfrom the first ceramic thermal barrier coatings, in that the porositiesof the first and second coatings are different, and the absolutedifference in the reduced or increased porosity being at least 2% whichcauses the respective coating thicknesses to be different, such that thedifference in the reduced and increased coating thickness is at least 50μm, and/or a single layer or a two layer nature of the coatings aredifferent; and wherein the new, second and/or the second turbine bladesor vanes are incorporated in the second gas turbine.
 2. The process asclaimed in claim 1, further comprising: removing a two-layer ceramicthermal barrier coating from the first turbine blades or vanes; and/orapplying a single-layer thermal barrier coating as a second ceramicthermal barrier coating to the second or new, second turbine blades orvanes.
 3. The process as claimed in claim 2, further comprising: thesingle-layer ceramic thermal barrier coating has a porosity of 18%±4%.4. The process as claimed in claim 1, further comprising: removing asingle-layer thermal barrier coating from the first turbine blades orvanes and/or applying a two-layer thermal barrier coating as a secondceramic thermal barrier coating to the second or new, second turbineblades or vanes.
 5. The process as claimed in claim 1, wherein a secondporosity of the second ceramic thermal barrier coating of the second orof the new, second turbine blades or vanes is elevated compared to afirst porosity of the thermal barrier coating of the first turbineblades or vanes.
 6. The process as claimed in claim 1, wherein a secondporosity of the second ceramic thermal barrier coating of the second orof the new, second turbine blades or vanes is lower as compared to thefirst porosity of the thermal barrier coating of the first turbineblades or vanes.
 7. The process as claimed in claim 1, furthercomprising replacing a thinner ceramic thermal barrier coating which isthe first ceramic thermal barrier coating by a thicker ceramic thermalbarrier coating which is the second ceramic thermal barrier coating ofthe second or of the new, second turbine blades or vanes; and thedifference in the thicknesses is at least +50 μm.
 8. The process asclaimed in claim 1, further comprising replacing a thicker ceramicthermal barrier coating which is the first ceramic thermal barriercoating by a thinner ceramic thermal barrier coating which is the secondceramic thermal barrier coating of the second or of the new, secondturbine blades or vanes; and the difference in the thickness being atleast −50 μm.
 9. The process as claimed in claim 1, further comprisingproducing the two-layer thermal barrier coating having a bottommostceramic layer having a porosity of 12%±4% and having an outer ceramiclayer having a porosity of 25%±4%, wherein an absolute difference in theporosities of the ceramic layers is at least 2%.
 10. The process asclaimed in claim 1, further comprising producing the bottom layer of thetwo-layer thermal barrier coating as thinner, than the top layer inwhich the total coating thickness of the two-layer thermal barriercoating is 500 μm to 800 μm.
 11. The process as claimed in claim 1,wherein partially stabilized zirconium oxide is used for both of thebottom ceramic layer and the top ceramic layer.
 12. The process asclaimed in claim 1, further comprising: using zirconium oxide for theceramic thermal barrier coating or the ceramic layers in a monoclinicproportion of the powder to be sprayed of less than 3%.
 13. The processas claimed in claim 11, wherein a tetragonal proportion of the ceramicthermal barrier has the greatest proportion in zirconium oxide.
 14. Theprocess as claimed in claim 12, further comprising performing a heattreatment for reducing the monoclinic proportion of the zirconium oxide,in a form of a powder to be sprayed.
 15. The process as claimed in claim1, further comprising spraying the bottom layer without polymer andspraying the top layer (10′) with polymer.
 16. The process as claimed inclaim 1, wherein the mean pore diameter of the top ceramic layer isgreater than the mean pore diameter of the bottom ceramic layer.
 17. Theprocess as claimed in claim 1, further comprising using the same powderwith the same composition and with the same grain size distribution forboth of the first or top and the second or bottom layer.
 18. The processas claimed in claim 1, further comprising using a different material forthe bottom ceramic layer than for the top ceramic layer.
 19. A gasturbine comprising turbine blades or vanes, wherein at least some of theblades or vanes have a single-layer ceramic thermal barrier coatingthereon with a porosity of 18%±4%.
 20. A gas turbine comprising turbineblades or vanes, wherein at least some of the blades or vanes have atwo-layer thermal barrier coating which comprises a bottommost ceramiclayer having a porosity of 12%±4% and a top ceramic layer having aporosity of 25%±4%, and the absolute difference in the porosities of theceramic layers is at least 2%.
 21. A gas turbine, comprising turbineblades or vanes in which the zirconium oxide of the ceramic coating orof the ceramic layer has a monoclinic proportion of less than 3%. 22.The gas turbine as claimed in claim 18, in which the bottom ceramiclayer comprises partially stabilized zirconium oxide and the outerceramic layer comprises partially stabilized zirconium oxide.
 23. Thegas turbine as claimed in claim 18, in which the outer layer of theceramic coating has a perovskite or pyrochlore structure, and the bottomceramic layer comprises zirconium oxide.
 24. The gas turbine as claimedin claim 18, in which the porosity of the bottommost ceramic layer hasbeen set by grain sizes of the powder to be sprayed, and the porosity ofthe outer ceramic layer has been set by particles of the powder to besprayed having smaller grain sizes with polymer.
 25. The gas turbine asclaimed in claim 20, in which the mean pore diameter of the top ceramiclayer is greater than the mean pore diameter of the bottom ceramiclayer, by at least 20 μm.
 26. A method for operating a gas turbinesystem, in which the nature of ceramic coatings of turbine blades orvanes of the gas turbine is altered, in particular as claimed inclaim
 1. 27. The method as claimed in claim 26, further comprisingstarting the gas turbine daily.
 28. The method as claimed in claim 26,further comprising continuously operating the gas turbine for severaldays.
 29. A method for operating a gas turbine system, using a gasturbine as claimed in claim
 20. 30. The process as claimed in claim 4,wherein only a two layer thermal barrier coating is applied to the newon second turbine blades or vanes.
 31. The process as claimed in claim18, further comprising using zirconium oxide for the bottom layer andusing a pyrochlore for the top layer.